Bio Aviation Technology Corp. System Definition Review

Team IV

Berger, Richard Daugherty, William Ksander, Jeff Lashkari, Dheer Malda, Jon Poulin, Christine Schreiner, Chris Voss, Bethany

AAE 451 March 23, 2006

Team IV, System Definition Review

Table of Contents

I. Executive Summary ------1 II. Introduction and Problem Definition ------2 III. Design Targets and Trade Studies ------3 Aircraft Sizing ------3 IV. Concept Selection------6 Initial Concepts------6 Pugh’s Method------12 Effects of Wing Location ------15 Ducted Fan Analysis------16 Selected Concept: Sodalis------17 V. Alternative Fuel ------18 Evaluation of Ethanol vs. Bio-diesel------20 VII. Design Constraint Analysis------23 VIII. Federal Aviation Regulations------28 IX. Conclusions ------30 References ------31 Appendix ------I Appendix A: QFD Matrix ------I Appendix B: Pugh’s Method, First Iteration ------II Appendix C: Acquisition Cost Model and Data------III

March 23, 2006 Team IV, System Definition Review

I. Executive Summary

Today’s world is facing depleting oil reserves as well as increasing public environmental concerns. These events cause a significant rise in fuel prices. Bio Aviation Technology Corporation (BAT Co) has determined that this trend creates a unique business opportunity. The company will develop an alternate fuel aircraft to satisfy customer aviation needs well beyond the affordable availability of petroleum based fuels.

Research of the current market has shown that the greatest profit potential resides within the small aircraft general aviation market. The aircraft being created is a single engine piston powered aircraft designed for the trainer market. Trade studies show that range and speed greatly effect the gross take off weight and therefore the overall cost of the aircraft. A greater understanding of the affects of various parameters help to design an aircraft suited for the targeted customers.

The Cessna 172 is used as a benchmark to compare with the BAT Co Sodalis aircraft. Through preliminary studies the designed aircraft is projected to have a longer range and a greater speed then the Cessna 172. The Sodalis will also have decreased internal and external noise, and create a greater thrust per horse power because of the use of a ducted fan instead of a conventional propeller.

What makes the Sodalis unique is the use of an alternate fuel. The new aircraft will be powered with bio-diesel. Bio-diesel is:

• Widely available • Performs similar to Jet-A fuel • Easy to produce • Requires minimal component conversion to utilize an existing engine

The bio-diesel specific fuel consumption (SFC) is slightly higher than an equivalent diesel aircraft engine. The estimated SFC is 0.45 and 0.36 respectively. The energy density and heat of combustion are both lower in bio-diesel and fuel weight is slightly higher than diesel. However, it is believed that the rising cost in petroleum based fuels due to a diminishing supply will ultimately result in lower operating costs.

Bio Aviation Technology Corporation engineers have made it the top priority to provide the customer with an aircraft to get them safely to any destination at an extremely competitive price as well as provide an excellent aircraft.

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II. Introduction and Problem Definition

One of the most important and concerning issues in the 21st century is the depleting oil resources. The net crude oil price has risen over the past 5 years and this trend is expected to continue as time progresses [1]. Petroleum-based fuel caters to almost all the transportation industries. One of the few industries left to successfully develop an alternate energy source is the aviation industry. There is also an increasing concern regarding environmental issues like depletion of the ozone layer, due to harmful emissions from petroleum based fuels. Bio Aviation Technology Corporation (BAT Co) envisions a business opportunity to service the aviation market by developing an alternate fuel trainer aircraft to replace the conventional piston driven aircraft, thereby providing a competitive alternative to the current market.

The BAT Co business strategy entails developing a general trainer aircraft design that can cater to a wide general aviation market. The primary target customers are hobbyists, small businesses, fixed based operators (FBO), flight schools and universities. Apart from the primary product, a general aviation aircraft, BAT Co anticipates a potential market for conversion kits to enable the use of alternative fuel technology in existing aircraft. Another aspect to the business case is providing a comprehensive in-house parts and maintenance service to the customer as well as maintenance personnel training.

The BAT Co aircraft will begin sales in 2010. By then, it is estimated that there will be an 11% decrease in petroleum-based aircraft from the prior year. BAT Co will supplement the 11% decrease in petroleum aircraft sales with the first year sales of the alternate fuel aircraft. The company will then begin the first year with 11% of the market and sell 124 planes. The BAT Co market share will then increase in years to come.

From 2011 until 2015, BAT Co will turn the petroleum market’s loss into a company gain plus 5%. An example is in 2011 when the petroleum aircraft sales are down 13%, BAT Co sales will increase to 18%. The total market share for the company in the year 2015 will be 40%. 2015 also marks the year in which it is decided that other manufacturers will begin to sell alternate fuel aircraft. BAT Co will continue to hold 40% of the market share and increase sales of aircraft by 5% annually until 2025. In 2026, the total number of aircraft sold is projected to be 670 per year. Based on General Aviation Manufactures Association (GAMA) in 2000, Cessna sold over 700 Cessna 172s and 182s, therefore 670 aircraft is a reasonable sales goal for BAT Co [2].

From 2026 until 2030 BAT Co will continue to sell 670 aircraft per year. Table 1 summarizes the market outlook for the company. Sales of alternate fuel aircraft by competing companies will increase. The total annual sales of the BAT Co aircraft and competitors’ alternate fuel aircraft are estimated to be 1675 in the year 2030. Also in that year, it is expected that for the first time no petroleum based aircraft will be manufactured. In the year 2001, 1791 petroleum-based general aviation aircraft were sold. By 2030, the 1675 total alternate fuel aircraft sold will again be over the 1600 mark which was about the total number of sales for piston powered aircraft in 2004.

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Table 1: Summary of BAT Co Market Share Total Aircraft Market Aircraft (Petroleum and Year Percentage sold Alternate Fuel) 2005 0% 0 1685 2010 11% 119 1119 2015 40% 297 737 2020 40% 524 1288 2025 40% 665 1675 2030 40% 670 1675

III. Design Targets and Trade Studies

A set of design targets was established through market research and a database of current aircraft. Desired values as well as upper and lower bounds were determined through the Quality Function Deployment (QFD) method, as shown in appendix A. Anything outside of these bounds would be unsuitable for the new alternative fuel aircraft. All of these bounds were developed from the aircraft database, since a new aircraft must meet or exceed current performance specifications to be marketed successfully.

A few of these targets have changed since the Systems Requirements Review in order to better accommodate the main target customer, the hobbyist. The changed targets include the range, useful load and cruise speed. All three of these design targets were reduced from the previous review based on a series of trade studies as well as additional market and aircraft research. The current set of design requirements are presented in Table 2.

Table 2: Revised Design Targets Minimum Desired Maximum Engineering Performance Performance Performance Characteristic Bound Bound Bound Units Acquisition Cost n.a. $172,500 $735,000 2005 USD Range 700 900 1,000 nm Takeoff Runway Length 3,000 1,200 n.a. ft Useful Load 700 800 1,300 lb Cruise Speed 125 175 200 kts Service Ceiling 11,000 15,000 n.a. ft Exterior Noise Level n.a. 60 70 dB Interior Noise Level n.a. 60 70 dB

Aircraft Sizing

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In order to perform trade studies involving the aircraft size and weight, a sizing code was developed. This code was developed by BAT Co and was based on the sizing approach presented in Aircraft Design: A Conceptual Approach by Daniel Raymer [3]. The inputs to the code include the crew and payload weight, cruise altitude and speed, aircraft range, specific fuel consumption, wing span and area, and the aircraft loiter time. The sizing code uses weight fractions for each mission segment to obtain the fuel weight fraction and take off gross weight of the aircraft. An iterative process was used to determine the aircraft gross take off weight (GTOW).

An acquisition cost model was developed from the aircraft database by using a regression to fit the data to a curve. Once the equation of this curve was known, it could be applied to the new aircraft to obtain an appropriate acquisition cost given the gross take off weight and cruise Mach number. The data used to build this model as well as the model itself can be found in appendix C. The sizing code can also be found in appendix C.

To ensure the accuracy of the sizing and cost model codes, these codes were validated by inputting the known parameters for the Cessna 172 and comparing the results with known Cessna performance and cost. The results are summarized in Table 3. It can be seen that all three predicted quantities are within ten percent of the actual Cessna data.

Table 3: Sizing Code and Cessna Comparison Cessna 172 Sizing prediction % Difference Take off weight [lbs] 2,450 2,425 1.02% Acquisition Cost [2005 USD] $172,500.00 $165,830.00 3.87% Horsepower [hp] 160 172.95 8.09%

Trade studies were then performed where the velocity, range, useful load and specific fuel consumption (SFC) were varied. The effects of these parameters on the gross take off weight and acquisition costs of the new airplane were then examined. Figure 1 below shows the gross take off weight vs. range with varying velocities. This study helped determined the feasibility of the desired performance specifications.

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Figure 1: Gross Take-Off Weight vs. Range

From Figure 1, it can be seen that as speed and range increase, the gross take off weight also increases. Since the cost is a direct correlation to the weight and speed the increased weight would cause the acquisition cost to increase. It can also be seen from the figure above that a change in velocity of 25 knots changes the gross take off weight by approximately 200 pounds. A change in range of 100 nautical miles only changes the weight by approximately 100 pounds. From this study, it was determined that a cruise speed of 250 knots was infeasible for this class of aircraft, and that design target was reduced to 175 knots.

While performing the trade studies, an important correlation was discovered. It was found that small changes in the useful load and SFC made large changes in the acquisition cost. By changing the useful load from 900 to 800 pounds, which is comparable to the Cessna 172, the acquisition cost is reduced by approximately $70,000. This cost reduction is extremely important for BAT Co, since acquisition cost is a large factor in the hobbyist and trainer market. Because cost is so important, BAT Co felt that the cost reduction was more important than the extra useful load capability.

It was also found that a specific fuel consumption change of 0.1 lb/hr/lb changes the acquisition cost by approximately $40,000. Since BAT Co’s aircraft is based on alternative fuel, it was important to gain an understanding of how changes to the specific fuel consumption change the aircraft weight and acquisition cost. The data from these studies can be seen in Table 4.

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Table 4: Effect of Useful Load and SCF on Gross Take-Off Weight and Cost

SFC Gross Take-off Acquisition Cost Useful Load [lbs] [lb/hr/lb] Weight [lbs] [2005 USD] 700 0.45 2,402 $170,100 800 0.45 2,679 $228,000 900 0.45 2,962 $298,500 700 0.35 2,213 $136,200 700 0.55 2,584 $206,900

All of the trade studies performed aided in providing a better understanding of what factors most affected the gross take off weight and acquisition cost of the new alternate fuel aircraft. Having an understanding of potential tradeoffs allowed BAT Co to make changes to the aircraft design targets, including speed, range and useful load, in order to best satisfy the potential customer. IV. Concept Selection

The BAT Co aircraft must be able to fulfill the customer’s requirements as well as be innovative in its design. Part of the innovation lies in the alternate fuel source and the remaining innovations must be incorporated into the airframe shape and layout. Target market provides the characteristics the aircraft must exhibit, as well as guidelines for the aircraft dimensions.

To find the best possible design, it is important to include a large number of choices in the selection pool. These possible designs have to be evaluated, based on the way they satisfy the customer’s requirements, and designs need to be eliminated until the best design remains. To accomplish this, BAT Co engineers have generated eight possible designs and evaluated them using Pugh’s method. The designs and the selection are described below.

Initial Concepts

During the concept generation, BAT Co engineers surveyed the current market to both determine the current design features and evaluated design trends as well as research aircraft design features that are not currently being utilized.

Based upon this research and individual design ideas, the engineers developed the following eight concept aircraft. The concepts range from traditional to unconventional designs and are detailed below.

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Figure 2: Concept I

The first concept, as seen in Figure 2, is a conventional design. The aircraft features a single- engine, tractor propeller, and low-wing. The T-tail is a design feature and also allows for a reduction in overall aircraft height through the endplate effect of the horizontal stabilizer [3]. The landing gear in this design utilizes a fixed, tricycle gear.

Figure 3: Concept II

The second concept is a less conventional than concept I. The aircraft features a canard configuration with a pusher ducted fan or shrouded propeller. In addition to the canard, this aircraft is equipped with a V-tail. The V-tail will be steep, since the canard already provides for a majority of the horizontal stabilizer area. This design can accommodate either fixed or retractable landing gear. In either case it will be a tricycle configuration.

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Figure 4: Concept III

Concept III is more conventional than the previous concept. The single-engine, tractor propeller, low wing aircraft is equipped with an H- or possibly a Tri-tail. This tail configuration allows for a reduction in aircraft height by splitting the total vertical surface area over multiple smaller surfaces [3]. In addition to this, the horizontal stabilizer area can be reduced due to the endplate effects of the vertical surfaces of the tail. This aircraft will also be equipped with a tricycle landing gear.

Figure 5: Concept IV

Concept IV is similar to the Adam Aircraft designs. The single-engine, pusher propeller, twin- boom configuration maximizes cabin space and reduces the structural weight and wetted area of the tail. The twin boom configuration allows for the use of a high-aspect ratio horizontal stabilizer, therefore reducing drag on the aircraft. The use of two vertical stabilizers also reduces the overall aircraft height. The major difference between a regular pusher propeller aircraft and concept IV is the fact that placing the propeller between the tail booms is an additional safety feature of the aircraft by guarding the propeller. Like previous designs, this aircraft will also be outfitted with a tricycle landing gear.

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Figure 6: Concept V

Concept V is the most unconventional of the proposed concepts. It features a pusher ducted fan, with a canard and vertical stabilizers and control surfaces located on the wingtips. This design takes advantage of the canard horizontal stabilizers and reduces overall aircraft height by splitting the total vertical control area into two separate surfaces. In addition, the location of the vertical surfaces gives the aircraft a very futuristic and sleek look. Concept V will also be fitted with a tricycle landing gear.

Figure 7: Concept VI

The sixth concept again exhibits more conventional features. It is a single-engine, tractor propeller, low wing aircraft. The Y-tail allows for a shallow V-tail on the top, because of the vertical stabilizer on the bottom, reducing aircraft height. In addition, the Y-tail will be a trademark feature, because it is not currently being used on any general aviation aircraft. The landing gear configuration will most likely be tricycle, although the tail configuration makes a tail-dragger configuration a possibility.

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Figure 8: Concept VII

The design goal of concept VII was to achieve a working balance between proven conventional and novel design features. This aircraft features a cantilevered high-wing, with a single ducted fan, in a pusher configuration. The conventional tail is used with this design to give it a more conventional appearance as well as keep the tail structural weight as low as possible. Like most of the other designs, this aircraft will be outfitted with a tricycle gear.

Figure 9: Concept VIII

As with concept VII, concept VIII aims to achieve a working balance between conventional and unconventional design features. However, this design incorporates a greater number of unconventional features to both achieve a more futuristic appearance and take advantage of their performance attributes. The canard will provide improved take-off and climb performance and the pusher ducted fan will improve low speed performance. Like concept VII, this aircraft will be outfitted with a tricycle landing gear as well.

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Figure 10: Concept IX

The last concept is similar to the first concept, with the exception of the wing placement. This aircraft features a high-wing with a T-tail and a tractor propeller. It will also be equipped with a fixed tricycle landing gear, like concept I.

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Pugh’s Method

To reduce the number of concepts and arrive at the best possible design, BAT Co engineers have employed Pugh’s method. This method is a way to evaluate a design by comparing its attributes to a set reference or datum design. This datum may be an existing, competing, design or one of the concepts. The evaluation is based on whether a concept satisfies a characteristic better than, equal to, or less than the datum. This scoring system clearly shows which parts of the design are already more than satisfying and which still have room for improvement, allowing for a concept to be qualitatively evaluated.

Specifically, Pugh’s method is carried out in several steps. After the concepts have been generated and the attributes to be compared have been selected, the first step is to run the comparison matrix. This means comparing every characteristic of every concept to the datum and assigning a score of better, same or less. Once the matrix is complete, each category is summed up, meaning that every concept will receive three scores, one equaling the number of better attributes, one equaling the number of attributes that perform the same and one for the number of attributes with lower performance. This is illustrated in the example Pugh’s matrix in Table 5.

Table 5: Example Pugh's Matrix Characteristics / Attributes Concept 1 Concept 2 Datum Attribute A + + Attribute B S + Attribute C S -

Number of “+” 1 2 Datum Number of “S” 2 0 Number of “-“ 0 1

After completing the matrix for the first time, each concept must be examined individually. The goal of this step in the evaluation is to improve the poorly performing attributes and enhance the other attributes. This may be accomplished by changing the features of a concept, combining it with other concepts and adding a completely new concept or eliminating it. Therefore, a new set of concepts will be created, which have to be evaluated in the same manner as the first ones. A new datum can be set based on the most advantageous concept.

The matrix is run again, the scores totaled, and the concepts evaluated again. This process is repeated until a single design consistently emerges as the top design.

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Concept Selection

For the new BAT Co aircraft, Pugh’s Method was used in order to reduce the number of possible designs. The concepts were evaluated based on the following criteria:

• Weight • Short Takeoff • Cruise Speed • Durable / Rugged • Stability • Short Manufacturing Time • Turnaround Time • Acquisition Cost • Exterior Noise • Appearance • Quiet Interior

During the first iteration, the initial concepts were compared to one of the main competitors of the new aircraft, the Cessna 172.

To ease the comparison, several aircraft features, such as tail configuration and wing location, were assigned as having positive or negative impacts on performance. Some of these values switched from positive to negative or vice versa, depending on the criterion being compared. For example, a high wing improves stability and is given a positive but increases turnaround time and is assigned a negative for that attribute. Other features, such as a T-tail, were generally assigned a negative value, because of the increased structural weight and adverse deep-stall characteristics.

To score the concept in a criterion, a positive feature balanced a negative, resulting in “same” and any additional positive or negative features shift the score to a multiple positive or negative. For each criterion, the concept was then compared with the datum and a score of better, same or worse performance was assigned. The first iteration comparison matrix is located in appendix B.

With the first run of the comparison matrix complete, the concepts were evaluated. It was discovered that after having improved the negative features of the concepts, several of them were now identical. This meant that the number of concepts could already be significantly reduced by simply not including multiple instances of the same concept.

The second comparison matrix utilized the same attributes for comparison as the first one; however, there were now only three concepts left in the selection pool. They were improved versions of concepts four and seven and a hybrid concept, formed through the combination and improvement of concepts II, V and VIII. Concepts I and IX were eliminated completely, because improving the designs by changing the tail configuration drastically changed the aircraft’s appearance and made them look too similar to airplanes currently in the market. The team believed that this would have an adverse effect on the marketability of the aircraft itself, therefore eliminated them from further consideration.

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The improvements made to concept IV entailed repositioning the wing, to create a more durable and rugged high-wing aircraft. Concept VII was improved by reconfiguring the tail to a cruciform configuration. This change both improved the appearance and reduced the duct construction complexity. The new hybrid concept turned out almost exactly like the improved concept seven, with the exception of the wing location; the Hybrid concept is a low-wing aircraft. The second comparison matrix is shown in Table 6, below.

Table 6: Second Comparison Matrix of Pugh's Method

' I' I' II II t t t p p p e e e m c c c u n n n t o o o a C C C D Cessna 172: High Wing; Convential Tail; forward prop High Wing; Boom Inverted V- Tail; Rear propeller High wing; ducted fan; Cruciform tail wing; Low ducted fan; tail Cruciform

Weight sss Short takeoff s - s Cruise Speed --s Durable/Rugged +++ Stability +++ Short manufacturing time + - s

Criteria Turnaround time --- Acquisition cost + - s Exterior Noise --s Appearance ---

Quiet interior --s DATUM

Number of + 422 Number of S 217 Number of - 582

For the second run of the comparison matrix, the engineers employed the same evaluation techniques used to complete the first run. In this case, the hybrid concept was used as the datum, because the team believed it to exhibit the most positive attributes and to gain an understanding of how the Cessna 172 compared to it. The run of the matrix resulted in the elimination of the twin-boom concept and showed that the Cessna 172 was inferior to the new reference design. The remaining concept, the improved concept seven, showed as much potential to be the best design as the datum, largely due to the fact that they only differ in wing location. Therefore, the team has decided to keep both designs and eliminate one after further study of the wing- placement effects. In addition to researching the wing placement, BAT Co must evaluate the advantages of a ducted fan and decide whether to move ahead or replace it with a conventional propeller.

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Effects of Wing Location

Pugh’s Method has aided in narrowing down the concepts, but two issues need more investigation before a final concept can be chosen. The first is whether or not to construct a high wing or low wing aircraft. Second, it must be determined whether the design concept should use a ducted fan or a propeller.

In order to make a decision on the wing placement, one must bear in mind the target market. Since this aircraft is intended for use primarily by less experienced pilots, the benefits of a high or low wing configuration to this group must be considered. A high wing aircraft will be inherently stable since the center of gravity lies below the wings. Stability is an important criterion to deliver for these customers. Also, ground visibility will be much better for a high wing than a low wing as the wing will not obstruct the pilot’s view. An added advantage to BAT Co is the fact that manufacturing a high wing is simpler because no dihedral needs to be built into the wing to gain a stability benefit.

Looking at disadvantages of the high wing configuration, one can see that visibility above the aircraft is limited and a heavier landing gear support will be required. Overhead visibility can be improved by using some kind of transparent building material for the upper part of the cockpit and even into the leading edge of the wing if necessary. The increased weight of the landing gear support is a trade off that must be made in order to reap the benefits of the high wing configuration.

There are some aerodynamic drag differences between a high and low wing configuration. Airflow over the top of the wing should be as undisturbed as possible to achieve maximum performance. On any low wing aircraft, the fuselage is placed directly in this flow while the high wing configuration places the fuselage beneath the wing in a flow that is less sensitive to disturbances. Due to the high wing’s location, less complex geometry is needed in order to minimize drag because of the favorable pressure gradients at that location [6].

Ground effect also needs to be considered when choosing the wing configuration. Ground effect generally impacts the aircraft approximately one wingspan above the ground. The closer the wing gets to the ground, the more pronounced the effect. Since low wings are closer to the runway than high wings, the ground effect is more prevalent in low wing aircraft. Due to this effect, low wing aircraft have the reputation for "floating" in the landing flare which may increase the required field length.

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Ducted Fan Analysis

Qualitatively, ducted fans appear to have much better performance than propellers over the operating speeds of the aircraft. The improved performance of a ducted fan for lower airspeeds includes greater thrust for a given horsepower. In fact, ducted fans can produce as much thrust as a propeller with a 40 percent greater diameter [7]. This increased thrust is especially valuable to reducing takeoff roll [8]. In a pusher configuration, a ducted fan has a stabilizing effect on the pitch and yaw moments much like the fletching on an arrow [6]. Again, any increase in stability for this aircraft is extremely valuable due to the trainer market BAT Co is targeting. The ducted fan is quieter than a propeller due to the enclosed fan. Noise damping materials can be also used in the duct for further noise reduction [7].

Another benefit of the duct is increased safety. The duct not only protects the fan from objects that might strike it, but also prevents people on the ground from being injured by the fan [7]. The precise values for increased performance of the ducted fan over the propeller are still being analyzed and must be weighed against the disadvantages of a ducted fan.

Ducted fans are generally heavier than propellers, but with the increased availability and knowledge of composite materials this weight increase can be minimized. Designing a duct that outperforms a propeller is a process that requires fine tuning of the duct and fan shapes. This precision along with the probable use of composite materials means manufacturing the ducted fan will be more expensive than manufacturing a propeller. The most important drawback of the ducted fan is that as the aircraft speed increases, the drag of the duct increases to the point at which any thrust benefit over the propeller is lost to this drag increase. BAT Co is working on a way to predict at which conditions this significant drag increase will occur.

The final disadvantage of the ducted fan is that maintaining it is more complex than maintaining a propeller because of the enclosed fan. BAT Co believes that the design concept employing a high wing configuration and a ducted fan will best serve the target customers, and efforts to obtain quantitative data of the benefits of this design are currently being made.

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Selected Concept: Sodalis

Based on the evaluation of the initial designs and the analysis of the wing placement, BAT Co engineers have decided to further develop the Sodalis, which is based on the improved version of concept VII. The Sodalis is a four-seat, single-engine, pusher ducted fan aircraft, aimed at the replacement of the Cessna 172 in the hobbyist market.

At this stage of the design, the Sodalis features a more spacious and quiet cabin than either the Cessna 172 or 182 due to the streamlined fuselage layout and the power plant location aft of the cabin. In addition, the fuselage offers ample room for bulkier luggage. The high-wing and large windows offer an excellent view of the ground for both the pilots and the passengers. The high- wing and the ducted fan combine to give the Sodalis improved short- and rough-field handling capabilities.

The new BAT Co aircraft also excels in the areas of safety and stability. It incorporates a ducted fan, which reduces the risk of injuries from a spinning propeller. In the safety area, the aircraft also features improved inherent stability due to the stabilizing effects of the high-wing and the pusher ducted fan. The aircraft also features excellent upward visibility for a high-wing aircraft, due to the large windshield. Shown in Figure 11 below is an illustration of the Sodalis.

Figure 11: Sodalis Illustrations

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V. Alternative Fuel

Initially, a variety of types of alternative fuels were considered. Among these were hydrogen, methanol, propane/kerosene, electric/solar, ethanol, bio-diesel and ethanol/bio-diesel blends. These fuels were compared using a variety of metrics. Comparison of energy densities can be seen in Figure 12. The numerical data is shown in Table 7. The specific fuel consumption (SFC) was taken into account during the fuel selection process, as well as logistics such as acquisition and storage.

14000

12000

10000

8000

6000

4000 Energy/Mass [kcal/kg] Energy/Mass 2000

0 Gasoline (l) Ethanol (l) Methane (l) #2 Diesel B100 AGE85 AVGas

Figure 12: Gravimetric Energy Density

Table 7: Numerical Data for Figure 12 Density Energy/Vol Energy/Mass Fuel Type [g/cm^3] [kcal/m^3] [kcal/kg] Gasoline (l) - 7.60E+06 1.06E+04 Ethanol (l) - 5.09E+06 6.44E+03 Methanol (l) - 3.80E+06 4.80E+03 Methane (l) - 8.59E+06 1.19E+04 Methane (g) (5000 psi) - 2.75E+06 1.19E+04 Hydrogen (l) - 2.03E+06 2.86E+04 Generic Fuels [7] Generic Fuels Hydrogen (g) (5000 psi) - 6.58E+05 2.86E+04 #2 Diesel 0.85 8.62E+06 1.01E+04 B100 0.88 7.88E+06 8.95E+03 B20 0.856 8.47E+06 9.90E+03 B2 0.851 8.61E+06 1.01E+04 AGE85 0.755 5.87E+06 7.78E+03 AVGas 0.715 7.49E+06 1.05E+04 Jet (wide-cut) 0.762 7.92E+06 1.04E+04

Conv [9] Diesel [8] Jet (kerosene) 0.81 8.37E+06 1.03E+04

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Although the government is pushing hydrogen based power for automobiles, the technology and distribution infrastructure does not exist yet. Predictions are that the infrastructure will not be mature when the Sodalis goes to market in 10 years. In addition, hydrogen fuels require either a large volume of space for fuel tanks which is infeasible for the scale of aircraft under consideration, or they require a heavy containment tank which would decrease the range and payload. Hydrogen fuel cell technology is overcoming some of these restrictions, but it was felt that the technology was not yet mature enough for application in the Sodalis.

Propane is widely available. However, it comes from a non-renewable feedstock, as does kerosene. Since BAT Co is looking to design an aircraft that runs on a completely renewable, environmentally friendly fuel, both propane and kerosene were ruled out.

Although considered initially, electric was ruled out due to the weight of the batteries and accessibility. The target market includes hobbyists who would be storing their aircraft at community airfields where power is not necessarily accessible. An alternative would be a removable battery pack that the operator could charge at home, which they would then have to transport safely to the airport. Battery technology is also not yet developed enough to deliver the amount of power a general aviation aircraft would require for longer flights while staying within a decent weight allotment. Also, the weight does not decrease through the course of the flight, which is a safety concern during landing. Considering solar power, the required wing area to accommodate solar cells would force a designed wingspan incompatible with most available general aviation hangars.

Ethanol was one of the top choices as a fuel. It has the advantage of already being type certified for use in the Cessna 172, the main competing aircraft [10]. Ethanol is easily renewable from a widely available feedstock. It works in off the shelf engines with major modification, and is becoming more widely available, both in pure ethanol and blended forms.

Methanol fuel has much of the same energy characteristics as ethanol. This can clearly be seen in figure 10 or in its numerical counterpart, table 6. Methanol is also more corrosive than ethanol, more dangerous to handle and transport as it is more volatile, and it is less beneficial to the environment.

Bio-diesel was also one of the top choices for fuel. It is completely renewable, and can be produced with relatively limited equipment in comparison to ethanol. Bio-diesel is widely available in most areas and requires only minor modifications to work in an existing engine. It also provides burn and energy characteristics similar to existing conventional fuels, as well as favorable lubrication characteristics that can help improve the performance and lifespan of an engine.

Although of comparable merit to bio-diesel and ethanol, ethanol/bio-diesel blends such as AGE- 85 were ruled out in favor of more conventional alternative fuels due to the lack of conclusive data on their characteristics.

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Evaluation of Ethanol vs. Bio-diesel

After taking a look at the wide variety of fuels available, BAT Co engineers chose to focus on pure ethanol and pure bio-diesel as the most viable options. These fuels were the most appropriate for the size of the Sodalis.

Ethanol is a proven fuel for aircraft. Supplemental type certifications have been established to convert general aviation aircraft such as the Cessna 172 to 100% ethanol. There is definitive data available for ethanol performance, and the use of ethanol for flight has been proven in many general aviation aircraft.

The conversion of existing hardware to allow the use of 100% ethanol is difficult and complex. Ethanol will electrolyze any non-anodized aluminum and steel parts, as well as react with most rubbers. Ethanol benefits from a much higher compression ratio than its conventional counterpart. To obtain this higher compression ratio, this may entail boring out the engine cylinders of an existing engine. The cylinders and piston heads would then need to be recoated with a non-reactive coating. Ethanol is also conductive, so any electrically driven parts such as the fuel pump would need to be replaced with more expensive, shielded versions. [11]

Also of concern is availability. Although ethanol, specifically E-85 is gaining in popularity, it is still rare. Currently there are 28 fuel stations in Indiana where E-85 is available, and only one listed is an airport. [12] Based on the current availability of E-85, BAT Co can not predict a wide spread availability of pure ethanol in the near future.

Along with availability is the ability to personally produce the fuel. Bio-diesel is relatively easy to produce on site, which BAT Co believes will be of interest to our customers. The production of ethanol requires a process similar to distilling liquor. This process is more dangerous and requires more equipment and more knowledge than the production of bio-diesel.

Converting a plane requires replacing almost the entire fuel system. Because BAT Co is interested in using a common off-the-shelf engine to decrease manufacturing costs as well as offering conversion kits as a part of the business plan, the complexity of this conversion is of concern. As an alternative, the upgrade required to use bio-diesel requires only the replacement of incompatible rubber components.

The main downside to using bio-diesel as a primary fuel source is that there are currently very few aircraft diesel engines available, and these are only beginning to be tested with bio-diesel fuels. There are currently no supplemental type certifications (STC) available for bio-diesel aircraft use, but it is believed that these will be available in the near future.

The Thielert Centurion 1.7 is one of the only in-production certified aircraft diesel engines. It is the engine being used as a baseline. The Centurion 1.7 has a STC available to retrofit it to the Cessna 172, however the upgrade costs around $40,000 to perform. [13]

March 23, 2006 Page 20 of 32 Team IV, System Definition Review

Another downside to bio-diesel is the fact that it gels at cold flight conditions. This is currently being overcome with fuel heaters and additives, but both these solutions can drive up the operating costs of the aircraft.

Bio-diesel is more widely available and easier to produce than ethanol. Bio-diesel also contains about 1/3 more energy than ethanol, which is important for a small, lightweight aircraft such as the one being considered. Because of the merits of bio-diesel over ethanol, BAT Co has chosen to focus on bio-diesel as the primary fuel source.

Because there is no current data on bio-diesel aircraft engines, BAT Co is scaling the performance specifications using known values from the automotive industry. In Figure 13, a graphical representation of the scaling of one of the more critical aspects of design, the SFC is shown. The average known SFCs for automotive diesel and bio-diesel power plants were examined and from those a percent difference was found. This number was then applied to the known SFC for the Centurion 1.7. [14] From this scaling, a theoretical SFC value of 0.45 for an aircraft bio-diesel engine was determined. The SFC is comparable to the Cessna 172, which has a SFC between 0.43 and 0.48.

0.43 0.50

% Δ SFC

Average automotive diesel Average automotive bio-diesel

0.36 0.45

Aircraft diesel Theoretical aircraft (Centurion 1.7) Bio-diesel

Figure 13: SFC Calculation Method

March 23, 2006 Page 21 of 32 Team IV, System Definition Review

Currently, bio-diesel costs more per gallon than diesel. The cost of a trip is increased by the fact that bio-diesel has a lower energy density and a higher SFC, and therefore requires more fuel. BAT Co believes, however, that as the supply of petroleum decreases, the cost of a gallon of diesel will increase. Additionally, as the demand for bio-diesel increases, the production and availability will increase, and the cost per gallon will decrease. Bio-diesel also produces less harmful emissions than diesel, as can be seen in Table 8. Based on the alternative fuel analysis, the Sodalis aircraft will be powered by a bio-diesel engine. [17]

Table 8: Environmental Merits of Bio-Diesel [15] Bio-Diesel Changes In Emissions (Compared to Diesel) Hydrocarbons -52.4% Carbon Monoxide -47.5% Nitrous Oxide -10.0% Carbon Dioxide 0.9%

March 23, 2006 Page 22 of 32 Team IV, System Definition Review

VII. Design Constraint Analysis

Having an understanding of the current design targets, aircraft concepts, and fuel research, more details regarding vehicle definition can be determined. The vehicle characteristics that most govern further vehicle design at this stage are takeoff roll-out, the ground-roll on landing, stall speed, and the maximum aircraft cruise speed.

Based upon the target design requirements as outlined previously, several comparison studies were performed. These studies allow for the determination of wing loading and power loading combinations that satisfy the different vehicle characteristics. From these studies, a constraint diagram was obtained. The results of the comparison are sets of intersecting curves. These curves bound a region of wing and power loading which satisfy vehicle capabilities, or a feasible design space.

A computer program was developed that was based upon the wing and power loading relationships as presented in Raymer [3]. These relationships require the input of certain vehicle characteristics such as wing area, aspect ratio, and mission segment weight ratios. Using this program, benchmarks were made using existing aircraft. These benchmarks allow for examination into the accuracy of the methodology. Specifically examined were the Cessna 172 and the Piper Warrior. Using manufacturer provided data when possible, and engineering approximations for those characteristics not readily available from the manufacturer, the program was validated. The results of the validation can be seen in figures Figure 14 and Figure 15 for the Cessna 172 and the Piper Warrior respectively.

Figure 14: Cessna 172 Constraint Analysis

March 23, 2006 Page 23 of 32 Team IV, System Definition Review

Figure 15: Piper Warrior Constraint Analysis

For a Cessna 172, the manufacturer published wing loading is 14.1 pounds per square foot. This point is denoted in Figure 14 along the maximum cruise curve. At this point, there is a horsepower to weight ratio of approximately 0.06527, resulting in a code predicted horsepower of 159.9. According to manufacturer specifications, the 172 has a 160.0 horsepower engine installed. The discrepancy between the constraint predicted code and the actual engine is very small. With this in mind, one will note that the point of wing loading as published does not fall within the design space in Figure 14. The landing and stall constraint lines limit the region of acceptable loading to 12.1 pounds per square foot. This is also found in the constraint coding for a Piper Warrior as shown in Figure 15. From this it is notable that the constraint coding provides a more conservative value of wing loading. At this point in the design process, a conservative estimate for wing loading is acceptable.

With an understanding of the accuracy of the constraint analysis predictions it was possible to construct a constraint curve for the current design target values. Vehicle parameters that have not yet been determined or governed by current design targets were based upon those on the Cessna 172. The use of these parameters based on the Cessna 172 is justifiable as it is the vehicle which has been identified as the current day primary competition. The inputs for the current design constraint analysis can be seen in Table 9. The maximum lift coefficient was based upon an un-flapped NACA 2412. At this point in the design there has been no selection of flaps or other high-lifting devices for landing or takeoff, thus the selection of a maximum lift coefficient based upon an un-flapped airfoil is justified. This results in a conservative estimation of power loading and at this point in the design process is acceptable.

March 23, 2006 Page 24 of 32 Team IV, System Definition Review

Table 9: Current Design Constraint Inputs Constraint Parameter Input Value Units Wing Span 36.1 FT Wing Area 174 SFT Oswald Efficiency 0.85 Propeller Efficiency 0.85 Stall Speed 61 KTS Maximum Cruise Speed 175 KTS Cruise Altitude 5000 FT Maximum Lift Coefficient 1.6 Takeoff Rollout Distance 1200 FT Landing Rollout Distance 1000 FT Landing to Takeoff Weight Ratio 0.8032 Induced Drag Coefficient 0.03

Using the parameters from Table 9, a constraint analysis was performed. The resulting constraint curves can be seen in figure 15. From the constraint curves, an allowable design space is bound by the maximum cruise speed and stall curves. The minimum acceptable wing loading occurs at a value of 17.35 pounds per square foot and a minimum acceptable horsepower of approximately 315. Notably, the power loading for the current design is higher than that of the Cessna 172 and Warrior. The relationships used to relate the power and wing loading are based on a pure propeller aircraft and the current design calls for the use of a ducted fan. The constraint curve as pictured in Figure 16 does not account for the benefits of a ducted fan, especially in regards to horsepower advantages. This is an area of continued research and examination, and at the current stage in the design process, the wing loading for design is based upon the minimum intersection point of the maximum speed and stall curve to minimize the required horsepower.

Figure 16: Current Design Constraint Analysis

March 23, 2006 Page 25 of 32 Team IV, System Definition Review

As was seen in Figure 16, maximum cruise speed was the most dominant constraint. Although these results do not incorporate the ducted fan advantages, it is desirable to see the effect on horsepower requirements based on differing maximum speed requirements. The maximum cruise speed was varied for several different values. The resulting maximum cruise speed curves can be seen in Figure 17. From this it is observable that the horsepower requirements drop significantly with maximum cruise speed. Obviously, because of the cost of maximum cruise speed, it is imperative to account for the advantages of the duct.

Figure 17: Current Design Maximum Cruise Speed Comparison Constraint Curves

Overall, from the vehicle parameters used initially for the constraint curves and the vehicle concept as outlined previously, a vehicle has begun to take shape. A dimensioned 3-view drawing of the Sodalis is presented in Figure 18.

March 23, 2006 Page 26 of 32 Team IV, System Definition Review

Figure 18: BAT Co Sodalis Dimensioned 3-View Drawing

Since this aircraft is sought to replace the Cessna 172, it is important to compare the Sodalis with the Cessna 172. The direct comparison can be seen in Table 10. Data presented in Table 10 has been determined based upon the analysis completed to date, including customer analysis, cost and weight modeling, fuel research, and constraint analysis.

Table 10: Comparison of the BAT Co Sodalis to the Cessna 172 Design Cessna 172 Units Engineering Characteristic Target Acquisition Cost $228,000 $172,500 2005 USD Range 900 687 nm Takeoff Runway Ground Roll 1,200 945 ft Useful Load 800 830 lb Cruise Speed 175 122 kts Service Ceiling 15,000 14,200 ft Exterior Noise Level 60 63 dB Interior Noise Level 60 92 dB Power to Weight Ratio 0.1179 0.0654 hp/lb Wing Loading 17.35 14.1 lb/ft2 Gross Take-off Weight 2679 2560 lb Wing Span 36 36 Ft Specific Fuel Consumption 0.45 0.42 lb/hr/lb

March 23, 2006 Page 27 of 32 Team IV, System Definition Review

VIII. Federal Aviation Regulations

Regulations become an important factor in aircraft design. All aircraft must pass FAA certification type tests for airworthiness as defined in chapter 1 of the Federal Aviation Regulations (FAR) guidebook. The FAR set a list of flight tests and design constraints that must be met in order to ensure the aircraft performs as intended. The federal regulations for the certification of fixed-wing, propeller-driven aircraft date back to 1926 and are codified in the Air Commerce Act. Then due to politics, changes in technology and aircraft accidents, the FAA issued new aircraft certification regulations in 1965 [18]. One of the major parts of the FAR is the type identification and certification. The FAR is divided into two main groups, small airplanes and transport airplanes. For small airplanes, FAR Part 23 defines the certification class of this aircraft design in Table 11.

Table 11: Federal Aviation Regulations Class A Certification Requirements Federal Aviation Regulations (FAR) Applicability Category: Various (Normal, Utility, Acrobatic, Agricultural) MTO (lb): < 6000 Number of Engines: 1 Engine Type: All Minimum Crew: Flight Crew: 1+ Cabin Attendants: None Maximum Number of Occupants: 10 Maximum Operating Altitude (ft): 25,000 Class: A

This categorization into Class A aircraft then further defines specifications, tests or constraints that the design must eventually meet. Many of the FAR flight tests are intended so that the manufacturer or designer can prove that their aircraft will in fact perform as predicted. However, other specifications, such as stall speed, structural loading factors, and balked take-off climb out rates become significant design constraints and must be considered accordingly. It is common to continuously cross-reference the specification design constraints with the FAR in order to ensure that requirements are met. Other FAR parts provide regulations for aircraft operating envelopes, propellers, noise requirements, and engine airworthiness. Table 12 is a list of the FAR Parts that are applicable to Class A aircraft design.

Table 12: Federal Aviation Regulations for Class A aircraft Design PART SUBPART DESCRIPTION 23 23.1 to 23.1589 Airworthiness Standards: Normal, Utility, Acrobatic, And Commuter Category Airplanes 33 33.1 to 33.99 Airworthiness Standards: Aircraft Engines 35 35.1 to 35.47 Airworthiness Standards: Propellers 36 36.1 to 36.1583 Noise Standards: Aircraft Type And Airworthiness Certification 91 91.1 to 91.1443 General Operating And Flight Rules

March 23, 2006 Page 28 of 32 Team IV, System Definition Review

Each of these FAR sections can be broken down into sub-sections. In part 23, most of the airworthiness factors are broken down into structural, dynamic, stability, take-off, landing, climb and safety constraints. These are often quantifiable values that are defined for each aircraft classification. Table 13 highlights the most important values for each section, and these most significantly affect the aircraft design.

Table 13: Most Notable Class A Federal Aviation Regulation Requirements FAR Designation Item Description/ Value 23.45 Reserve Fuel (night, VFR/IFR) 45 min 23.49 Maximum Stall Speed (Vso) <61 kts 23.337 Structural Load Factors (n) +3.8, -1.590 23.303 Factor of Safety 1.5 23.25 Passenger Weight Assumption (lb) 170 23.73 Landing Approach Speed (Va) >1.3Vso 23.51 Minimum Takeoff Speed (Vto=Vr) >1.1Vso 23.51 Climb-out over 50-ft height (Vcl) >1.2Vso 23.233 Safe-landing crosswind velocity (Vxw) >0.2Vso

As can be seen, the 45 min. fuel reserve and stall speed become significant design constraints to the aircraft design, as they affect large design values such as maximum lift coefficient (CLmax), range and fuel weight.

March 23, 2006 Page 29 of 32 Team IV, System Definition Review

IX. Conclusions

Bio Aviation Technology Corp. has selected an aircraft that will meet the needs of the target customers. The aircraft is primarily designed to be an excellent trainer, and entry level aircraft. Some features of the plane include a high wing design, cruciform tail and a ducted fan. The ducted fan will help to reduce the internal and external noise, as well as provide a greater thrust per horsepower than a conventional propeller, and reduce the required takeoff roll. The weight and drag increase caused by the ducted fan can be off-set by the use of composite materials and optimized duct geometry.

Compared to a Cessna 172, the BAT Co aircraft is projected to be faster, quieter and have a longer range while still being competitively priced. Bio-diesel will be used to fuel the aircraft and has a comparable SFC to current diesel aircraft engines. It also has lower emissions and is projected to cost less than petroleum based fuels as the supply of petroleum decreases.

Constraint analysis has provided insight into the power and wing loading requirement, called for in the current design, and has also highlighted the costs and tradeoffs that can be made within the design parameters. Additionally, it has highlighted the need for an improved model, which incorporates the performance of a ducted fan.

Additional analysis is necessary to further develop the design. Most of the design challenges emerge as a result of the pusher ducted fan propulsion system. The ducted fan places additional weight into the tail of the aircraft and the engine is located at a rather far aft position. The solution to this challenge is two fold. First, an engine has to be found that is light and must be located as far forward as possible. Second, the tail and duct structure must be as light as possible to reduce the moment due to their weight.

Work continues in the areas of researching manufacturing costs, composite material analysis, design specifications and ducted fan analysis. The BAT Co engineers are confident that the new alternate fuel aircraft will satisfy the needs of target customers and will continue to solidify the design of the Sodalis.

March 23, 2006 Page 30 of 32 Team IV, System Definition Review

References

1. Campbell, C.J. Peak Oil. Presentation at the Technical University of Clausthal [electronic version]. (2000). Retrieved January 21, 2006 from http://www.peakoil.org/

2. 2000-2004 Shipment Report, 2005 Third Quarter Shipment Report. General Aviation Manufacturing Association [Electronic Version]. (2005). Retrieved on January 20, 2006 from http://www.gama.aero/resources/statistics/shipments.php

3. Raymer, Daniel P. (1999). Aircraft Design: A Conceptual Approach. 3rd Edition. American Institute of Aeronautics and Astronautics (AIAA).

4. Davisson, B. High Wing vs. Low Wing. Plane and Pilot magazine (February 2002) [Electronic Version] Retrieved on February 16, 2006 from http://www.planeandpilotmag.com/content/pastissues/2002/feb/highwing.html

5. Stoen, H. High Wing/Low Wing. (April 2002) [Electronic Version] Retrieved on February 16, 2006 from http://stoenworks.com/High%20wing,%20Low%20wing.html

6. McClellan, J. Mac. “Left Seat.” FLYING April 2006: 13-18.

7. Appendix C: Propulsion. Virginia Tech. Pg 1-18 [Electronic Version] Retrieved on February 15, 2006 from filebox.vt.edu/eng/aero/ga2002/WordDocs/AppC_Propulsion.doc

8. Gundry, Matt. Ducted Fan Design for a Cozy Mark IV [Electronic Version]. Retrieved on February 15, 2006 from http://faa-engineers.com/~mjgundry/ductedfan/

9. Roan, V. P, Fletcher J. A Comparison of Several Alternative Fuels for Fuel-Cell Vehicle Applications [Electronic Version]. American Institute of Aeronautics and Astronautics, 1994. Retrieved on February 6, 2006. AIAA-94-3796-CP.

10. National Bio-diesel Board. Energy Content. Retrieved February 2, 2006 from http://www.biodiesel.org/pdf_files/fuelfactsheets/BTU_Content_Final_Oct2005.pdf

11. Chevron Products Company. Aviation Fuels Technical Review. Chevron U.S.A Inc, 2006. Retrieved February 2, 2006 from http://www.chevron.com/products/prodserv/fuels/bulletin/aviationfuel/2_at_fuel_perf.shtm

12. FAR STC SE8707SW. Federal Aviation Administration (2006). Retrieved on February 2, 2006 from http://www.airweb.faa.gov/Regulatory_and_Guidance_Library/rgSTC.nsf/0/C034D02C3F

13. Weins, William et. al. The New Silverado: An Ethanol (E85) Conversion by the University of Nebraska-Lincoln. Society of Automotive Engineers, 2000.

March 23, 2006 Page 31 of 32 Team IV, System Definition Review

14. National Ethanol Vehicle Collation. Indiana E-85 Refueling Stations. Retrieved February 14, 2006 from http://www.e85fuel.com/database/locations.php?state=inIndiana.

15. Kohl, Ronald. Throw out the Textbooks, Diesel Airplanes are Here. Machine Design, 2003. Retrieved February 2, 2006 from http://www.machinedesign.com/ASP/strArticleID/55900/strSite/MDSite/viewSelectedArticle .asp

16. Thielert AG. Centurion 1.7. Retrieved February 14, 2006 from http://www.centurion- engines.com/c17/c17_data.htm.

17. Vern, Hofman. Biodiesel Fuel. Retrieved on February 14, 2006 from http://www.ext.nodak.edu/extpubs/ageng/machine/ae1240w.htm

18. Kimberlin, Ralph D., (2003). Flight Testing of Fixed-Wing Aircraft. American Institute of Aeronautics and Astronautics (AIAA).

19. Code of Federal Regulations (e-CFR) [Electronic Version]. Retrieved on March 16 2006 from http://ecfr.gpoaccess.go

March 23, 2006 Page 32 of 32 Team IV, System Definition Review

Appendix

Appendix A: QFD Matrix

p i P

s s e C

34 25 13

e i u Q

032 055 9 024 1

n i

l i

c i

v r e S

99 1 1 3044 3 0022 0055 1

m u N

3 30352 9 9 9 0 00

i r

e t x E

333332 9 000043 00 11 000044 0 0 1

i t

u q c A

9 003 9 0 000055 19 9 1 1

i t

s a *Take-off decibal level (6500 m from roll-out) F

19 0 0 9 03

i t

r o h S

00 0 19 00 0 000053 9 9 1 9199

t i l

b a t S

0 33 9 30330 3

s i u r C

91 1 0300

l y a P

3 30000 11

a t

r o h S

9 19 919

an r

g n o L 11 9191

184 210 239 211 280 199 119 413 235 279 184 221

6.63% 7.57% 8.62% 7.61% 10.09% 7.17% 4.29% 14.89% 8.47% 10.06% 6.63% 7.97%

h g i We 7 6 53 3 0000000 0 00000 9 00000002 0 000055 000000 4 00003 23 12110003 3 15 0 0 1033333 0 120000000 10123333330 0 0 0 Units nm ft lbs kts Yes/No days min USD '05 dB # ft dB Cessna 182 845 795 1213 145 Yes 40 35 326150 *70 4 18,100 75 Target Value 1000 1200 1300 250 Yes 30 30 172500 *60 4 15,000 60 bsolute Piper Seneca 819 1143 1367 190 Yes 45 40 735000 *64 6 18,500 70 Importance Relative Importance A Fast range Long runways on most Land/Take-off to fly Easy bag and golf skis Take hanger a standard in Fit Safety & available quick Maintenance time Turnaround to maintain Cheap Comfortable Edge Cutting Durable Fast deliverable time deliverable Fast to buy Cheap good Looks Performance Time Costs Other

March 23, 2006 Page I Team IV, System Definition Review

Appendix B: Pugh’s Method, First Iteration

prop forward DATUM

Convential Tail; Tail; Convential

High Wing; Wing; High

Cessna 172: 172: Cessna D

forward prop forward

t t

p Wing; T-tail; T-tail; Wing;

n - -

s s s o

Traditional; High High Traditional; C

Change vertical tail vertical Change Fan

tail; Rear Ducted Ducted Rear tail;

I I

Canard; Vertical Vertical Canard;

p - e s

High Wing; Wing; High C

Ugly; High wing High Ugly; Rear Ducted Fan Ducted Rear

I I

t t

p Conventional tail; tail; Conventional

n -- --

s+ ++s ++s o

High Wing; Wing; High C

Propeller

Forward Forward

Low Wing; Y-tail; Y-tail; Wing; Low

I - - -- - s s

Concepts V

t t

n o

concepts 2) concepts canard

(becomes 'fixed' 'fixed' (becomes Ducted fan; fan; Ducted

add another tail tail another add Winglet tail; Rear Rear tail; Winglet

e s s +

Take off Winglet tail; tail; Winglet off Take Low Wing; Wing; Low C

location; low wing low location; Rear propeller Rear

V Boom Tail; Prop Prop Tail; Boom Inverted V-Tail; V-Tail; Inverted

c Low Wing; Boom Boom Wing; Low ------ss + s s+

n o

Prop

I Tail; Forward Forward Tail;

t t

c - --

n s s s Low Wing; H- Wing; Low o

surfaces; V-tail surfaces; Fan; Canard Fan;

t t

p redundant control control redundant Rear Ducted Ducted Rear

n -- - -

s o Low Wing; V-tail; V-tail; Wing; Low C

Better then Datum then Better Datum as Same Worse then Datum

forward prop forward

I Wing; T-tail; T-tail; Wing;

c - - --

Traditional; Low Low Traditional; s s sss +s+s+ s ++s s s +s+ss ss+ ++s +s+s+ s ++s 151 3 6827 3 3243 2 5 2 5 5 6 0 4 3 2 9 3 3 2 n + +++ + + + s s s o C LEGEND: Weight Stability Number of - Number of + Number of S Number Appearance Short takeoff Short Quiet interior Cruise Speed Cruise Exterior Noise Exterior Durable/Rugged Acquisition cost Acquisition Turnaround time Turnaround Negative Features: Features: Negative

Short manufacturing time manufacturing Short Criteria

March 23, 2006 Page II Team IV, System Definition Review

Appendix C: Acquisition Cost Model and Data

Appendix C.1: Acquisition Cost Data

Min 1.048% Max 25.286% Cost Difference % Average 11.801% Calculated Calculated se i ft ($) Cost ru Mach Mach C Number @ 10,000 @ 10,000 Speed Cruise (knots) (lb) Max Max Weight Weight Takeoff Takeoff $ = 1.6031e-4Wto^2.6941 * * Mcr^0.1440 PiperPiperPiperPiper Warrior Mirage 2,440 6XT Saratoga 4,340 115 3,600 213.00 3,600 0.1802 175 0.33 161 $195,300 $1,100,000 0.2742 0.2522 $167,400.30 $863,147.44 $572,400 $415,700 14.286% $507,080.48 21.532% $501,010.09 11.412% 20.522% Cirrus SR22 3,400 185 0.2898 $349,750 $438,187.65 25.286% Cessna Skyhawk 172182 2,450 3,100 Skyhawk Cessna 122 Skylane Cessna 145 0.1911 0.2272 $172,500 $326,150 $170,692.52 $329,883.77 1.048% 1.145% Cessna StationairCessna 3,600 142 0.2225 $448,160 $492,051.66 9.794% Manufacturer Model Aircraft Diamond Aircraft Star 2,535 241.2 0.3779 $204,000 $206,423.87 1.188%

March 23, 2006 Page III Team IV, System Definition Review

Appendix C.2: Sizing Code

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % William Daugherty % Jeff Ksander % Team IV - BATCO % Sizing Code – Handler Function %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % handler.m %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% clear all close all clc

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Set Simulation Variables %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

Wcrew = 300 % Crew Weight, LBS Wpayload = 600+30 % Passenger Weight, LBS Altitude = 8000 % Cruise Altitude, FT Velocity = 122 % Cruise Speed, KTS Range = 700 % Cruise Range, NM E = 45 % Loiter Time, MIN

C = 0.45/60/60 % Prop Specific Fuel Consumption, Raymer PG 23 b = 36 % Wing Span, FT S = 174 % Wing Area, FT^2 e = 0.8 % Oswald Effiency Factor

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Iterate on Values % Velocity in KTS Below %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

[hp,W0,cost] = sizing(Wcrew,Wpayload,Altitude,Velocity,Range,C,b,S,E,e);

% p = 1; % v = [60:1:180]; % for m = v(1):v(length(v)) % [hp(p),W0(p),cost(p)] = sizing(Wcrew,Wpayload,Altitude,m,Range,C,b,S,E,e); % p = p + 1; % end % % %Velocity = 125 kts % p = 1; % r = [400:1:1000];

March 23, 2006 Page IV Team IV, System Definition Review

% for n = r(1):r(length(r)) % [hp2(p),W02(p),cost2(p)] = sizing(Wcrew,Wpayload,Altitude,Velocity,n,C,b,S,E,e); % p = p + 1; % end % % %Velocity = 150 kts % Velocity = 150; % p = 1; % r = [400:1:1000]; % for n = r(1):r(length(r)) % [hp3(p),W03(p),cost3(p)] = sizing(Wcrew,Wpayload,Altitude,Velocity,n,C,b,S,E,e); % p = p + 1; % end % % %Velocity = 175 kts % Velocity = 175; % p = 1; % r = [400:1:1000]; % for n = r(1):r(length(r)) % [hp4(p),W04(p),cost4(p)] = sizing(Wcrew,Wpayload,Altitude,Velocity,n,C,b,S,E,e); % p = p + 1; % end % % %Velocity = 200 kts % Velocity = 200; % p = 1; % r = [400:1:1000]; % for n = r(1):r(length(r)) % [hp5(p),W05(p),cost5(p)] = sizing(Wcrew,Wpayload,Altitude,Velocity,n,C,b,S,E,e); % p = p + 1; % end

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Format Plots %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% plot(v,W0) xlabel('Velocity (KTS)') ylabel('Gross Take Off Weight (LBS)') title('Gross Take Off Weight vs. Velocity') figure(2) plot(r,W02,r,W03,'r',r,W04,'g',r,W05,'m') xlabel('Range (NM)') ylabel('Gross Take Off Weight (LBS)') title('Gross Take Off Weight vs. Range') legend('Velocity = 125 KTS', 'Velocity = 150 KTS', 'Velocity = 175 KTS', 'Velocity = 200 KTS')

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Add Parameters to Plots Automatically (I Don't Wanna Do This Crap % Everytime %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

March 23, 2006 Page V Team IV, System Definition Review

figure(1) useful = Wcrew + Wpayload; load_str = ['Useful Load = ',num2str(useful)]; range_str = ['Cruise Range = ',num2str(Range)]; alt_str = ['Cruise Altitude = ',num2str(Altitude)]; smat = [load_str, ', ', range_str, ', ', alt_str]; ax = axis; xloc = 1.05*ax(1); yloc = 1.04*ax(3); text(xloc,yloc,smat) figure(2) useful = Wcrew + Wpayload; load_str = ['Useful Load = ',num2str(useful)]; % velocity_str = ['Cruise Velocity = ',num2str(Velocity)]; alt_str = ['Cruise Altitude = ',num2str(Altitude)]; smat = [load_str, ', ', alt_str]; ax = axis; xloc = 1.05*ax(1); yloc = 1.04*ax(3); text(xloc,yloc,smat)

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

Sizing Code function [hp,W0,cost,p] = sizing(Wcrew, Wpayload, Altitude, Velocity, Range,... C,Span,S,E,e)

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % William Daugherty % Jeff Ksander % Team IV - BATCO % Sizing Code %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Inputs: % Wcrew Crew Weight, LBS % Wpayload Crew Payload, LBS % Altitude Cruise Altitude FT % Velocity Cruise Velocity KTS % Range Cruise Range NM % C Specific Fuel 1/SEC % Span Wing Span FT % S Wing Area FT^2

March 23, 2006 Page VI Team IV, System Definition Review

% E Loiter Time MIN % e Wing Efficiency Factor %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Outputs: % hp Required Engine Horse Power, HP % W0 Gross Take Off Weight, LBS % p Iterations to Converge, %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % sizing.m %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % W0 Model Coefficient Definitions % Raymer, PG 116 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% a = -0.25; b = 1.18; C1 = -0.20; C2 = 0.08; C3 = 0.05; C4 = -0.05; C5 = 0.27;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Cd0 Induced Drag Model Coefficient Definitions % http://pr.erau.edu/~gallyt/ae302/oldhomeworks.html % aa = -2.046 Fixed Gear % aa = -2.222 Retractable Gear, Aluminum Skin % aa = -2.301 Retractable Gear, Composite Skin %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% aa = -2.046;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Horse Power Model % Based on GA Trends %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

[DENSITY,SSD] = stp(Altitude); Mcr = Velocity*1.687809857/SSD; hp = 0.0390*Mcr^0.3924*(Wpayload+Wcrew)^1.3240;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Conversion Factors

March 23, 2006 Page VII Team IV, System Definition Review

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% kts2fts = 1.687809857; nm2ft = 6076.115485564; min2sec = 60;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%% % Initial Calculations and Initial Weight "Guess" %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%

A = (Span^2)/S; % Aspect Ratio W0 = 50000; % Initial Weight "Guess" temp = 10000; % Transfer Variable for W0 p = 1; % Iteration Count Initialize ep = 1; % Error Initialize while ep > 1e-4

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Variable Transfers and Cd0 Calculation Raymer

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

temp = W0; CDo = 10^(aa + 1.089 +0.515*log10(W0))/S; % Induced Drag Model W0 = (W0);

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Take Off & Roll Out Weight Fraction Raymer

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

W1W0 = 0.97;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Climb Fraction - Uses Mach Number - Raymer PG 117 % Calls stp.m Atmospheric Tables for Speed of Sound at Final Altitude

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

[DENSITY2,SSD2] = stp(Altitude); W2W1 = 1.0065 - 0.0325*Velocity*kts2fts/SSD2;

March 23, 2006 Page VIII Team IV, System Definition Review

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Cruise - Simplified Fraction Based on Raymer EQ 6.12 and ETA for % Props EQ 3.9.

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

W = W0*W2W1*W1W0; W3W2 = exp(-Range*nm2ft*C*hp/(Velocity*kts2fts*W)); %%

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Loiter - Simplified Fraction Based on Raymer EQ 6.15 and ETA for % Props EQ 3.9

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

W = W0*W3W2*W2W1*W1W0; W4W3 = exp(-E*min2sec*C*hp/W);

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Desecent Raymer

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

W5W4 = 0.990;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Landing Raymer

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

W6W5 = 0.992;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Calculating Total Fractional Values % WeW0, Raymer PG 116, Table 6.2

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

WxW0 = W1W0*W2W1*W3W2*W4W3*W5W4*W6W5; Wx = WxW0*W0; WfW0 = 1.06*(1 - WxW0);

March 23, 2006 Page IX Team IV, System Definition Review

num = -b*(C1-C3+C4)*(Wpayload+Wcrew)*(Velocity^C5)*((W0/S)^C4)*... (A^C2)*(W0^C1)*((hp/W0)^C3); den = W0*(b*(Velocity^C5)*((W0/S)^C4)*(A^C2)*(W0^C1)*((hp/W0)^C3)... +a+WfW0-1)^2;

fpW0 = num/den; fW0 = a + b*(W0^C1)*(A^C2)*((hp/W0)^C3)*((W0/S)^C4)*(Velocity*kts2fts)^C5;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Determine W0 and Find Error

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

Wo = W0; W0 = W0 - (fW0/fpW0); ep = abs((W0-Wo)/W0); p = p+1;

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%% % Kills Script if Failing to Converge - Graphs Will Look Like CRAP % When Convergence Fails! Nothing Worthwhile Comes From More Than % 1000 Iterations

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%

if p > 500000 fprintf('\n!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!') fprintf('\nSOLUTION CONVERGENCE FAILED') fprintf('\n!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!!\n\n') break break end end

cost = (1.6031e-4)*W0^(2.6941)*Mcr^(0.1440); return

March 23, 2006 Page X